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Dual-Mode Free-Jet Combustor Dual-Mode Free-Jet Combustor Charles J. Trefny and Vance F. Dippold III [email protected] NASA Glenn Research Center Cleveland, Ohio USA Shaye Yungster Ohio Aerospace Institute Cleveland, Ohio USA ABSTRACT The dual-mode free-jet combustor concept is described. It was introduced in 2010 as a wide operating-range propulsion device using a novel supersonic free-jet combustion process. The unique feature of the free-jet combustor is supersonic combustion in an unconfined free-jet that traverses a larger subsonic combustion chamber to a variable throat area nozzle. During this mode of operation, the propulsive stream is not in contact with the combustor walls and equilibrates to the combustion chamber pressure. To a first order, thermodynamic efficiency is similar to that of a traditional scramjet under the assumption of constant-pressure combustion. Qualitatively, a number of possible benefits to this approach are as follows. The need for fuel staging is eliminated since the cross-sectional area distribution required for supersonic combustion is accommodated aerodynamically without regard for wall pressure gradients and boundary-layer separation. The unconstrained nature of the free-jet allows for consideration of a detonative combustion process that is untenable in a walled combustor. Heat loads, especially localized effects of shock wave / boundary-layer interactions, are reduced making possible the use of hydrocarbon fuels to higher flight Mach numbers. The initial motivation for this scheme however, was that the combustion chamber could be used for robust, subsonic combustion at low flight Mach numbers. At the desired flight condition, transition to free-jet mode would be effected by increasing the nozzle throat area and inducing separation at the diffuser inlet. Preliminary two-dimensional axisymmetric calculations with ethylene fuel and equilibrium chemistry are presented and discussed. They indicate feasibility of the unconfined supersonic combustion process and reveal shock and viscous losses unique to the free-jet concept. It was shown that variation of the nozzle throat area could be used to modify the free-jet shock structure through variation of the pressure in the recirculation zone surrounding the jet. Shocks were also initiated locally within the jet by combustion which began immediately at the inflow plane due to the equilibrium chemistry assumption. Performance and heat load assessments are described. Follow-on work that refines the initial equilibrium results to include the effects of finite-rate chemistry, and non- uniform fuel-air inflow profiles to more accurately assess ignition characteristics and combustion efficiency is presented. These calculations were carried out at a Mach 8 flight condition with ethylene fuel. V-gutter flameholders were used to initiate combustion at the combustor inflow station. Various fuel-air ratio profiles were imposed a short distance upstream of this plane to simulate upstream fuel injection. The effect of these profiles on thrust, wall heat flux and solution stability are presented. An unexpected result was an unsteady, periodic solution caused by intermittent ignition for some cases with a combustible mixture near the free-jet boundary. These results are included along with the recourse employed to stabilize the flowfield. Results with a smaller combustor diameter, relevant to a Mach 8 point design with reduced heat load and frontal area are presented. Also included are results at Mach 6 and 10 flight conditions for the same geometry. This was done to assess the feasibility of a fixed-geometry flowpath over this Mach number range. Plans for higher fidelity three-dimensional calculations with discrete fuel injection are outlined. Keywords: Airbreathing Propulsion; Supersonic Combustion Ramjet NOMENCLATURE M Mach number T Temperature A Cross-sectional area P Pressure R Radius V Velocity x Axial distance ṁ Mass rate of flow Symbols Skewness parameter in formula for Gaussian fuel profile Fuel-air equivalence ratio Subscripts 0 Freestream condition 1 Station 1 (x = 0), inlet throat and free-jet combustor inflow plane 8 Station 8, free-jet combustor nozzle throat A Air EQ Equilibrium T Total (stagnation) condition max Maximum 1.0 INTRODUCTION The potential for high speed and long range has driven aircraft designers to consider airbreathing propulsion since the dawn of high speed flight. The requirement for ever higher flight Mach number spurred development of ramjet propulsion. As the hypersonic flight regime was being explored, issues with the subsonic combustion ramjet cycle and implementation became apparent. These included materials limitations due to the severe stagnation conditions encountered in the combustion chamber, prohibitive momentum losses in the compression process, and impractical variable geometry requirements for the inlet and nozzle. It was recognized by early pioneers in high speed airbreathing propulsion1,2,3 that these problems could be relieved in a flowpath designed for supersonic combustion. Figure 1 is a diagram from reference 1 that shows the general layout of a supersonic combustion ramjet. Figure 1. – Diagram of a supersonic combustion ramjet from reference 1. In general, the cross-sectional area of the supersonic combustor increases in the downstream direction to avoid thermal choking and excessive pressure gradients. Processes that govern performance include inlet momentum losses, Rayleigh losses due to heat addition, heat loss to the combustor walls, skin friction, and chemical non- equilibrium. Other factors that must be considered include separation of boundary-layers due to adverse pressure gradients, intense local heating at re-attachment points and shock impingements, and fuel staging or variable geometry to accommodate the variation of combustion area ratio required with changes in freestream stagnation enthalpy. Airframe integration is also critical, as in using the vehicle forebody for compression and aft-body for expansion of the propulsive stream. The supersonic combustion ramjet or “scramjet” as it became known, cannot generate thrust at low flight Mach numbers, thus requiring some other means of acceleration to its operating condition. To this end, Curran and Stull4 introduced the “Dual Mode Supersonic Combustion Ramjet Engine” in a 1972 US Patent that proposed operation in a thermally-choked, subsonic combustion mode at low flight Mach numbers. A diagram from the patent appears in figure 2 where it can be seen that fuel is introduced upstream in supersonic combustion mode at high flight Mach numbers, and downstream in a larger cross-sectional area at low flight Mach number where the heat of combustion is sufficient to form a thermal throat and back-pressure the system. Figure 2. – “Dual Mode Supersonic Combustion Ramjet Engine” from reference 4. The cross-sectional area at which the thermal throat must form, increases as flight Mach number decreases, unless the fuel-to-air ratio is reduced. For a given duct, this effect determines the minimum flight Mach number for thermally-choked operation. At Mach 3, the required thermal throat area approaches that of the inlet capture area and it could be surmised that combustion extend into the nozzle expansion region. The primary technical challenges in practical application of the dual-mode scramjet scheme are modulation of the thermal throat location, fuel distribution, and ignition and flame-holding in the large cross-section. Any in-stream devices must be retractable or expendable so as not to inhibit supersonic combustion operation. The free-jet combustor concept was introduced in 20105 for application to a wide operating range propulsion system. An alternative to the dual-mode scheme described above, it evolved from a reversal of thinking if you will; a ramjet that operates in a supersonic combustion mode, instead of a scramjet with a thermal throat that forces subsonic combustion. The unique feature of the free-jet combustor pictured in figure 3a, is supersonic combustion in an unconfined free-jet that traverses a larger subsonic combustion chamber to a variable area nozzle throat. During this mode of operation, the propulsive stream is not in contact with the combustor walls, and equilibrates to the surrounding combustion chamber pressure. Qualitatively, a number of possible benefits to this approach are as follows. The combustion process within the jet is augmented by shock waves or potentially oblique detonation waves without regard for the cross-sectional area constraint imposed by walls. Shock wave / boundary-layer interactions, causing localized intense heating and combustor-inlet interaction are eliminated. Inlet flow uniformity requirements are relaxed. Variation in the combustor cross-sectional area distribution, required with flight condition and throttle setting changes, is accommodated aerodynamically by the free-jet without the need for fuel staging. Only the nozzle throat area, through which the jet must issue, is varied. Introduction of a secondary flow to the combustion chamber surrounding the jet may be used to reduce wall temperature and increase the combustion chamber pressure, thereby providing additional aerodynamic contraction. To a first order, thermodynamic efficiency is similar to that of a “traditional” scramjet under the assumption of constant-pressure combustion. The dual-mode aspect of this device is that the combustion chamber would operate efficiently as a subsonic combustion ramjet to low flight Mach number as pictured in figure 3b. Fuel is injected
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